HQ 2.5/9 AIRFOIL (hq259-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: HQ 2.5/9 AIRFOIL (hq259-il) Reynolds number: 500,000 Max Cl/Cd: 88.43 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hq259-il-500000-n5.txt Download as CSV file: xf-hq259-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: HQ 2.5/9 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.500 -0.4000 0.08360 0.08143 -0.0350 1.0000 0.0077
-8.250 -0.4012 0.08015 0.07802 -0.0362 1.0000 0.0077
-8.000 -0.4065 0.07645 0.07437 -0.0372 1.0000 0.0078
-7.750 -0.4172 0.07321 0.07119 -0.0373 1.0000 0.0079
-7.500 -0.4052 0.06654 0.06454 -0.0472 0.9893 0.0079
-7.250 -0.3845 0.05814 0.05603 -0.0605 0.9803 0.0080
-7.000 -0.3633 0.05114 0.04887 -0.0694 0.9730 0.0082
-6.750 -0.3427 0.04500 0.04251 -0.0750 0.9631 0.0081
-6.500 -0.3199 0.04179 0.03912 -0.0777 0.9550 0.0063
-6.250 -0.3009 0.03571 0.03270 -0.0808 0.9439 0.0055
-6.000 -0.2857 0.02480 0.02101 -0.0811 0.9308 0.0039
-5.750 -0.2621 0.02238 0.01820 -0.0809 0.9219 0.0037
-5.500 -0.2382 0.02031 0.01580 -0.0807 0.9135 0.0037
-5.000 -0.1891 0.01655 0.01135 -0.0799 0.8979 0.0036
-4.750 -0.1638 0.01488 0.00936 -0.0794 0.8910 0.0036
-4.500 -0.1381 0.01360 0.00784 -0.0790 0.8833 0.0036
-4.250 -0.1128 0.01219 0.00621 -0.0784 0.8764 0.0037
-4.000 -0.0874 0.01108 0.00493 -0.0779 0.8690 0.0039
-3.750 -0.0613 0.01029 0.00398 -0.0776 0.8622 0.0044
-3.500 -0.0345 0.00983 0.00343 -0.0773 0.8552 0.0051
-3.250 -0.0073 0.00951 0.00298 -0.0772 0.8486 0.0058
-3.000 0.0200 0.00927 0.00267 -0.0770 0.8414 0.0070
-2.750 0.0475 0.00902 0.00231 -0.0769 0.8344 0.0088
-2.500 0.0750 0.00885 0.00206 -0.0767 0.8267 0.0108
-2.250 0.1024 0.00860 0.00185 -0.0766 0.8182 0.0288
-2.000 0.1295 0.00838 0.00169 -0.0765 0.8095 0.0621
-1.750 0.1567 0.00812 0.00153 -0.0764 0.8005 0.1087
-1.500 0.1828 0.00744 0.00138 -0.0766 0.7922 0.2787
-1.250 0.2071 0.00657 0.00134 -0.0765 0.7845 0.5450
-1.000 0.2343 0.00647 0.00133 -0.0763 0.7755 0.5914
-0.750 0.2611 0.00640 0.00133 -0.0760 0.7653 0.6372
-0.500 0.2874 0.00634 0.00134 -0.0755 0.7534 0.6856
-0.250 0.3141 0.00634 0.00134 -0.0752 0.7397 0.7085
0.000 0.3411 0.00636 0.00132 -0.0749 0.7249 0.7220
0.250 0.3684 0.00640 0.00131 -0.0747 0.7117 0.7317
0.500 0.3957 0.00644 0.00131 -0.0746 0.6994 0.7415
0.750 0.4229 0.00648 0.00132 -0.0744 0.6863 0.7512
1.000 0.4498 0.00653 0.00134 -0.0741 0.6704 0.7611
1.250 0.4765 0.00661 0.00136 -0.0738 0.6527 0.7717
1.750 0.5295 0.00676 0.00146 -0.0732 0.6157 0.7951
2.000 0.5555 0.00686 0.00152 -0.0728 0.5931 0.8082
2.250 0.5808 0.00699 0.00159 -0.0723 0.5634 0.8228
2.500 0.6051 0.00718 0.00169 -0.0716 0.5262 0.8397
2.750 0.6289 0.00737 0.00180 -0.0708 0.4898 0.8609
3.000 0.6519 0.00751 0.00196 -0.0698 0.4570 0.8965
3.250 0.6853 0.00775 0.00211 -0.0711 0.4153 1.0000
3.500 0.7105 0.00809 0.00230 -0.0708 0.3805 1.0000
3.750 0.7352 0.00846 0.00251 -0.0704 0.3412 1.0000
4.000 0.7597 0.00887 0.00274 -0.0700 0.3013 1.0000
4.250 0.7845 0.00924 0.00301 -0.0696 0.2720 1.0000
4.500 0.8097 0.00958 0.00326 -0.0693 0.2479 1.0000
4.750 0.8351 0.00988 0.00350 -0.0690 0.2292 1.0000
5.000 0.8606 0.01017 0.00376 -0.0688 0.2117 1.0000
5.250 0.8841 0.01068 0.00407 -0.0682 0.1739 1.0000
5.500 0.9058 0.01137 0.00447 -0.0675 0.1241 1.0000
5.750 0.9297 0.01183 0.00486 -0.0670 0.1005 1.0000
6.000 0.9519 0.01247 0.00531 -0.0663 0.0669 1.0000
6.250 0.9736 0.01317 0.00584 -0.0656 0.0377 1.0000
6.500 0.9922 0.01428 0.00671 -0.0643 0.0039 1.0000
6.750 1.0157 0.01477 0.00727 -0.0636 0.0029 1.0000
7.000 1.0390 0.01528 0.00786 -0.0629 0.0026 1.0000
7.250 1.0617 0.01585 0.00853 -0.0622 0.0023 1.0000
7.500 1.0837 0.01649 0.00929 -0.0613 0.0022 1.0000
7.750 1.1049 0.01720 0.01012 -0.0603 0.0021 1.0000
8.000 1.1248 0.01807 0.01116 -0.0591 0.0021 1.0000
8.250 1.1432 0.01905 0.01228 -0.0578 0.0020 1.0000
8.500 1.1605 0.02009 0.01345 -0.0563 0.0020 1.0000
8.750 1.1760 0.02128 0.01478 -0.0546 0.0020 1.0000
9.000 1.1894 0.02263 0.01628 -0.0526 0.0020 1.0000
9.250 1.2026 0.02394 0.01773 -0.0506 0.0020 1.0000
9.500 1.2141 0.02538 0.01932 -0.0484 0.0020 1.0000
9.750 1.2231 0.02688 0.02097 -0.0459 0.0020 1.0000
10.000 1.2290 0.02854 0.02280 -0.0431 0.0020 1.0000
10.250 1.2346 0.03017 0.02460 -0.0403 0.0020 1.0000
10.500 1.2383 0.03202 0.02663 -0.0376 0.0020 1.0000
10.750 1.2368 0.03449 0.02932 -0.0347 0.0019 1.0000
11.000 1.2323 0.03731 0.03238 -0.0320 0.0019 1.0000
11.250 1.2278 0.04011 0.03539 -0.0297 0.0019 1.0000
11.500 1.2340 0.04140 0.03686 -0.0285 0.0018 1.0000
11.750 1.2369 0.04324 0.03886 -0.0274 0.0017 1.0000
12.000 1.2349 0.04584 0.04164 -0.0264 0.0017 1.0000
12.250 1.2304 0.04891 0.04488 -0.0259 0.0016 1.0000
12.500 1.2189 0.05314 0.04932 -0.0258 0.0016 1.0000
12.750 1.2140 0.05668 0.05303 -0.0263 0.0016 1.0000
13.000 1.2038 0.06133 0.05786 -0.0276 0.0015 1.0000
13.250 1.1898 0.06697 0.06368 -0.0298 0.0015 1.0000
13.500 1.1693 0.07435 0.07127 -0.0333 0.0016 1.0000
13.750 1.1576 0.08076 0.07784 -0.0369 0.0016 1.0000
14.000 1.1424 0.08851 0.08575 -0.0416 0.0015 1.0000
14.250 1.1271 0.09687 0.09426 -0.0467 0.0015 1.0000
14.500 1.0995 0.10874 0.10630 -0.0539 0.0017 1.0000
14.750 1.0866 0.11751 0.11518 -0.0591 0.0016 1.0000
15.000 1.0617 0.12965 0.12745 -0.0661 0.0017 1.0000
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Polar data table (+)
Polar graphs
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